The present disclosure relates to a gas turbine engine and, more particularly, to a combustor liner having film cooling circuits.
Gas turbine engines, such as those that power modem commercial and military aircraft, include a fan section to propel the aircraft, compressor section to pressurize a supply of air from the fan section, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust.
The combustor section may have an annular wall having inner and outer shells that support respective inner and outer heat shielding liners. The liners may be comprised of a plurality of floating heat shields or panels that together define an annular combustion chamber. An annular cooling plenum is defined between the respective shells and liners for supplying cooling air to an opposite hot side of the panels through a plurality of strategically placed film cooling holes. The film cooling holes are generally orientated to create a protective blanket, or, air film over the hot side of the panels, thereby protecting the panels from the hot combustion gases in the chamber.
Unfortunately, placing and/or distributing known film cooling holes across the panel to achieve an even temperature distribution is difficult. This difficulty is further aggravated due to structural obstacles and/or panel features that disrupt the formation of the cooling air film. Uneven temperature distributions or panel hotspots create thermal mechanical stresses that lead to cracking and a shortened panel lifespan.